1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine airfoil with trailing edge cooling channels.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine, especially an industrial gas turbine engine, includes a turbine section with multiple stages of turbine blades and stator guide vanes to convert the energy from a hot gas flow into mechanical energy to drive the rotor shaft. The efficiency of the engine can be increased by passing a higher gas flow temperature into the turbine. However, the highest temperature that the turbine can be exposed to is related to the material characteristics of the vanes and blades in the first stage. The higher the inlet temperature to the turbine, the higher will be the engine efficiency.
In order to allow for higher gas flow temperatures into the turbine, the turbine airfoils include complex internal cooling circuits to provide cooling for the airfoils. The engine efficiency is also increased by passing less cooling air through the airfoils for cooling. Since the cooling air used in the turbine airfoils is typically pressurized cooling air from the compressor of the engine, using less bleed off air from the compressor will also increase the engine efficiency.
A turbine rotor blade must be designed to not only have adequate cooling, but also be capable of withstanding the high centrifugal forces that develop on the blade from the rotation during operation. Also, the turbine rotor blades are subject to high temperatures that lower the material strength of the blades and can lead to creep problems from long exposure to strain. Erosion is also a problem in turbine airfoils if hot spots develop on portions of the airfoil that is not adequately cooled. Thus, it is desirable to provide for a turbine airfoil such as a turbine rotor blade with a minimum amount of material to reduce weight, and to provide for a maximum amount of cooling using a minimum amount of cooling air.
It is known in the art of turbine airfoil cooling that cooling efficiency can be improved by a reduction of the cooling channel wall thickness. However, for a low cooling flow design, as the airfoil wall thickness is reduced the internal cooling channel cross sectional flow area will increase. This will reduce the internal flow Mach number and through flow velocity, and thus reduce the cooling flow channel internal heat transfer coefficient as well as the channel convective performance.
U.S. Pat. No. 7,189,060 issued to Liang (the same inventor of the present application) on Mar. 13, 2007 and entitled COOLING SYSTEM INCLUDING MINI CHANNELS WITHIN A TURBINE BLADE OF A TURBINE ENGINE discloses a turbine blade with mini channels formed within the cooling channels along the blade spanwise direction of the serpentine flow cooling circuit. The channels are formed by ribs that have the same length throughout the channel from near the platform to near the tip. The mini channels of the present invention are formed in the trailing edge region of the blade in which the width of the blade decreases. The mini channels in the trailing edge of the blade of the present invention have different structure than the mini channels in the earlier Liang patent.
It is therefore an object of the present invention to provide for a turbine airfoil with a thin wall convection cooling channel along the trailing edge of the airfoil in order to improve the cooling of the trailing edge region.
It is another object of the present invention to provide for a turbine airfoil with a trailing edge cooling channel that will increase the cooling effectiveness without increasing the internal cooling channel air flow area so that the cooling effectiveness is increased.